JP6861827B2 - Adaptive machining of cooling turbine blades - Google Patents

Description

Cross-reference with related applications This application is incorporated herein by reference in its entirety, US provisional application No. 62/445956, filed January 13, 2017. Request priority of .62 / 445,956).

Background 1. The present invention relates generally to the manufacturing of turbine blades, in particular the process of adaptive machining of cast turbine blades with internal cooling passages.

2. Description of Related Techniques Gas turbine blades are usually manufactured by casting, especially investment casting. Cooled turbine blades have one or more internal cooling passages formed using the core during the investment casting process. The investment casting process imposes certain restrictions, especially on important wing features such as outer wall thickness, trailing edge thickness and shape. For example, as schematically shown in FIG. 1, during the casting process, for example, differences in solidification / shrinkage of metal parts can cause deformation and / or displacement (indicated by dashed lines). The example shown in FIG. 1 shows the core deformation in the form of twist or rotation in the case of the front edge cooling passage LE and the trailing edge cooling passage TE and the core displacement in the case of the chord center cooling passage MC. Deformation of the core can lead to changes in the shape and / or position of the cooling passages, which can deviate the wall thickness of the outer wall of the cast turbine blade from the nominal or target wall thickness of the turbine blade.

Casting restrictions as described above have some implications for the size and weight of the components. New generation gas turbine engines tend to have increased size turbine blades to achieve higher loads. The required blade geometry with thin blades can be difficult to manufacture by investment casting due to such process limitations. Historically, such casting restrictions with any blade size and shape have limited the design options available.

Summary Briefly, aspects of the invention provide techniques for adaptive machining of blades that may overcome certain casting process limitations, especially those involving core deformation and / or displacement.

According to the first aspect of the invention, there is provided a method of machining a blade section of a turbine blade or vane manufactured by a casting process. The wing section has an outer wall that defines the interior of the wing with one or more internal cooling passages. The method comprises receiving design data for the wing section, including nominal outer wing shape and nominal wall thickness data. The method further comprises generating a machined path by determining the target outer wing shape. The target outer wing shape is generated by adapting the nominal outer wing shape so that the nominal wall thickness is maintained at all points on the outer wall around one or more internal cooling passages in the subsequently machined wing section. Will be done. The method then involves machining the outer surface of the wing section manufactured by the casting process according to the generated machining path to remove excess material to follow the generated target outer wing shape.

According to a second aspect of the invention, there is provided a CAD module for generating machining path data for adaptively machining a blade section of a turbine blade or vane manufactured by a casting process. The wing section has an outer wall that defines the interior of the wing with one or more internal cooling passages. The CAD module is configured to receive design data for the wing section, including nominal outer wing shape and nominal wall thickness data. The CAD module is further configured to generate machining path data by determining the target outer blade shape. The CAD module targets the outer wing by adapting the nominal outer wing shape so that the nominal wall thickness is maintained at all points on the outer wall around one or more internal cooling passages in the wing section that will continue to be machined. It is configured to generate a shape. Machining path data provides information for machining the outer surface of the wing section manufactured by the casting process to remove excess material to follow the generated target outer wing shape.

The invention is shown in more detail with reference to the drawings. The drawings show a preferred configuration and do not limit the scope of the invention.

FIG. 6 is a schematic view of core deformation or displacement in an investment casting process for manufacturing turbine blades. FIG. 5 is a perspective view of a cast turbine blade including a blade section in which aspects of the invention may be practiced. FIG. 3 is a cross-sectional view taken along the cross section III-III in FIG. It is a schematic diagram which shows the composition of the point which represents the value of the nominal wall thickness around the measured position of the internal cooling passage in a wing section. It is the schematic which shows the best fit alignment of the nominal outer wing shape with respect to the said point which represents the value of a nominal wall thickness. FIG. 6 is a schematic showing a target outer wing shape that matches the final outer surface of the machined wing section. It is a schematic diagram which shows the system for adaptively machining a cast blade section by one aspect of this invention.

Detailed Description In the following detailed description of the preferred embodiments, the accompanying drawings forming a portion thereof are referred to, which, by way of example, are specific embodiments in which the invention may be practiced, but not as limited. The form of is shown. It should be understood that other embodiments may be used and modifications may be made without departing from the ideas and scope of the invention.

Embodiments of the present invention are shown in connection with turbine blades, typically large span blades that can be used in the low pressure turbine stages of gas turbine engines. It should be noted that aspects of the invention may be applicable to other turbine components having blade sections, such as rotating blades or fixed vanes in high or low pressure turbine stages.

Here, with reference to FIG. 2, a turbine blade 10 that may be manufactured by a casting process, such as an investment casting process, is shown. The cast turbine blade 10 has a blade section 12 extending radially outwardly from the platform 14 in a span direction with respect to a rotation axis (not shown). The blade 10 further has a root portion 16 extending radially inward from the platform 14 and configured to attach the blade 10 to a rotor disk (not shown). With reference to FIGS. 1 and 2 together, the cast wing section 12 is formed from an outer wall 18 that generally defines the interior of the hollow wing. The outer wall 18 has a generally concave positive pressure surface 20 and a generally convex negative pressure surface 22 connected at the front edge 24 and the trailing edge 26. The interior of the wing has one or more internal cooling passages 28 for radial flow of cooling fluid. The internal cooling passage 28 may be defined between the internal partition walls 30. The outer wall 18 has an outer surface 18a configured to face the high temperature gas path and an inner surface 18b facing the internal cooling passage 28.

The internal cooling passage 28 is formed by the casting core during the investment casting process. As mentioned above, during the casting process, for example, due to differences in solidification or shrinkage of metal parts, the core may undergo deformation (eg, rolling, rotation) and / or displacement. Deformation of the core can lead to changes in the shape and / or position of the internal cooling passage 28, which can deviate the wall thickness of the outer wall 18 from its intended thickness. Aspects of the invention at least solve the above-mentioned problems associated with core deformation and / or displacement.

According to embodiments of the present invention, the final shape of the wing section may be formed by adaptively post-machining the outside of the wing section (ie, the outer surface 18a of the outer wall 18) beyond the casting limits. .. As described herein with reference to FIGS. 3-6, the method for adaptive post-machining of cast wing sections includes wing section with nominal outer wing shape 40 _N and nominal wall thickness _TN data. receive design data relating to 12, to generate a machining path by determining a target outer airfoil shape 40 _T, the target outer airfoil shape 40 _T is the nominal wall thickness T _N is one in wing section to be machined later _{Alternatively, it is generated by adapting a nominal outer wing shape 40 N} so that it is maintained at all points on the outer wall 18 around the plurality of internal cooling passages 28, and is extra to follow the generated target outer wing shape 40 _T. The outer surface 18a of the wing section 12 manufactured by the casting process is machined according to the machining path in order to remove the material. The target outer wing shape 40 _T has been adapted to compensate for core shifts (deformations and / or displacements) during the casting process and takes into account the prioritized considerations of the following criteria in the order stated: Generated based on: 1) nominal wall thickness of outer wall 18 around the internal cooling passage 28 and 2) nominal wing outer shape.

In the first premachining step after the casting process, three-dimensional (3-D) measurements are made to determine the outer shape of the individual cast blade sections. The 3-D measurement is, for example, a tactile coordinate measuring machine probing, or laser scanning or photogrammetry, any combination thereof, or another other to obtain 3-D geometric data on the outer shape of the cast wing section. It may be carried out by the measurement technique of. The measured outer shape indicated by 3-D surface 40 _A in FIG. 4 corresponds to the outer surface 18a of the cast wing section 12 shown in FIG.

The next step involves making a cooling passage position and shape measurement for the internal cooling passage 28 with respect to the measured outer shape 40 _{A of the cast blade section 12.} Cooling passage position and shape measurement, as shown in FIG. 3, by performing the actual wall thickness measured at a plurality of points along the outer wall 18 of the cast airfoil sections 12 (shown as T _A) It may be done. The actual wall thickness is measured, have been uniformly denoted as T _A for simplicity, it should be noted that there may change at different points in the outer wall 12. Wall thickness measurements may be made using ultrasound, x-rays, calculated tomography, eddy currents or any other known technique. For example, in the case of measurement using an ultrasonic wall thickness T _A is a single signal transmitter / probe to a point on the outer surface 18a of the outer wall 18 of the wing section 12 is disposed, the outer wall 18 of the strongest echo signals are received It may be measured by determining the distance to one point on the inner surface 18b of the. By measuring the wall thickness values at a sufficiently large number of points along the axial (chong direction) and radial range of the outer wall 18, the 3-D geometry of the cooling passage is 28 m (as shown in FIG. 4). Shape and position) may be determined for the measured outer shape 40 _A of the cast wing section.

Further referring to FIG. 4, in a subsequent step, the point 42 is configured around the measured position 28 m of the internal cooling passage, which represents the value of _{the nominal wall thickness (TN) obtained from the design data.} That is, the points 42 are constructed at a distance equal _{to the nominal or design wall thickness TN} from each point on the periphery of the measured shape 28 m of the internal cooling passage. Point 42 may be formed along the radial span of the cooling passage. For simplicity, the nominal thickness is shown uniformly as _TN. One of ordinary skill in the art will recognize that the nominal thickness value can vary at different points around the internal cooling passage, both in the radial and axial directions (the chord direction).

Then, as shown in FIG. 5, the 3-D nominal outer wing shape 40 _N (obtained from the design data) is repeatedly best fit to match the point 42 representing the value of the nominal wall thickness _TN. The operation is performed. For an ideal casting process, all points 42 representing nominal wall thickness values are located on the _{nominal outer blade shape 40 N.} In the illustrated example, due to angular orientation and relative displacement of the casting core during the casting process, at least some of the points 42 deviate from the _{nominal outer blade shape 40 N after best fit alignment.} ..

Next, as shown in FIG. 6, a target outer wing shape 40 _T is generated by adapting the nominal outer wing shape 40 _{N after best fit alignment.} As shown in FIG. 6, a point representing the value of the nominal wall thickness deviating from _{the nominal outer wing shape 40 N} after the best fit alignment (ie, a point located inside or outside _{the nominal outer wing shape 40 N).} Are shown as 42a, whereas these points representing the nominal thickness values located on _{the nominal outer wing shape 40 N (or within the specified tolerance range) after the best fit alignment are 42b.} It is shown as. The target outer wing shape 40 _T is a 3-D shape generated by adjusting the 3-D nominal outer wing shape 40 _N , thereby deviating from the best fit alignment of the nominal outer wing shape 40 _{N 42a.} Is now located on the target outer wing shape 40 _T. Therefore, the target outer wing shape 40 _T coincides with all points 42a and 42b representing nominal wall thickness values, as shown in FIG. As mentioned above, the target outer wing shape 40 _T is to the prioritized criteria for adaptation: the nominal wall thickness ( _TN ) and the nominal outer wing shape (40 _N ) obtained from the design data. Obtained based on.

The above steps for generating the target outer wing shape 40 _T may be performed via computer-aided design (CAD) as described below. In the illustrated embodiment, the CAD module has a target outer wing shape 40 _T so that the target outer wing shape 40 _{T does not extend beyond the measured outer shape 40 A} of the cast wing section 12. May be adapted to limit.

Machining path data may be generated based on the target outer blade shape 40 _T. The machining path data provides information for machining the outer surface of the cast wing section corresponding to _{the measured shape 40 A} in order to remove excess material to follow the generated target outer wing shape 40 _T. Prescribe. Based on the generated machining data, the outer surface of the outer wall may be machined, for example, by grinding or milling. However, the outer wall machining may be performed by other means including, but not limited to, electrolytic machining (ECM) and electrical discharge machining (EDM) in particular.

The nominal wall thickness in the above-described embodiment, the casting process, (machining before) the actual wall thickness in the cast airfoil section 12 (T _A) is the corresponding obtained from design data of each of the wing sections It may be configured to guarantee that it is greater than the value of ( _TN). For post-machining of turbine blades or vanes in any turbine row, the machining of each blade section may be adapted to simultaneously adapt to the shape of the outer blade surface and the internal cooling passages. This creates a specific machining path for machining each wing section of the row of blades or vanes. Since the core deformation varies from individual blade to individual blade, the generation of machining paths and the execution of machining may be specifically adapted to each turbine blade.

Another aspect of the invention relates to an automated system for adaptive post-machining of cast wing sections. As shown in FIG. 7, such a system 50 may include a sensor module 52. As described above, the sensor module 52 measures the shape and position of the cooling passage by making a 3-D measurement of the outer shape of the cast wing section and measuring the actual wall thickness value of the cast wing section. Includes sensors for. The system 50 may have, for example, memory means 54 containing design data in the form of a 3-D model or CAD model of turbine blades or vanes. System 50 further receives measurement data 62 from sensor module 52 and design data 64 (eg, nominal wall thickness value, nominal outer wing shape) from memory 54 to generate machining path data 66 according to the method described above. ) Has a CAD module configured to receive. The CAD module may be a subcomponent for a computer-assisted design package. The machining path data 66 generated by the CAD module may include a numerical control (NC) program. The system 50 further includes a machining apparatus for machining the outer surface of the cast turbine blades based on the machining data 66. The CAD module may automatically set up, check and adapt the NC program for each cast turbine blade. It will be appreciated that CAD modules may be specified in the computer code and may be used to operate the computer to perform the methods described above. That is, a method and article that implements a computer code suitable for use in operating a computer to perform a method is an independently identifiable aspect of one invention concept.

The above embodiments, including adaptive machining of thin blades, may overcome casting process limitations, which allows non-castable geometry to be produced, eg, have thinner blades, taper. Allows the manufacture of wings with thinner trailing edges with no or small taper. The thinner blade outer wall can significantly reduce the centrifugal tensile load on the rotating turbine blades, especially in the low pressure turbine stage. The illustrated embodiment also allows for a more cost-effective manufacturing method as compared to reducing the wall thickness by casting process optimization. Another advantage is the possibility of reducing casting process tolerances and / or increasing casting wall thickness, which increases casting yield and thus reduces casting costs.

Although certain embodiments have been described in detail, one of ordinary skill in the art will recognize that various modifications and alternatives to these details can be developed in light of the overall disclosure. Accordingly, the particular sequences disclosed are intended to be exemplary only and not limiting with respect to the scope of the invention to which the full scope of the appended claims and all and all equivalents thereof should be given. ing.

Claims (8)

A method of machining a blade section (12) of a turbine blade or vane manufactured by a casting process, wherein the blade section (12) defines a blade interior having one or more internal cooling passages (28). The method has an outer wall (18).
Received design data for said wing section (12), including nominal outer wing shape (40 _N ) and nominal wall thickness ( _{TN) data.}
Generates a machining path by determining a target outer airfoil shape (40 _T), the target external airfoil shape (40 _T) is the nominal wall thickness (T _N) is one in wing section to be machined later Or generated by adapting the _{nominal outer wing shape (40 N} ) so that it is maintained at all points on the outer wall (18) around the plurality of internal cooling passages (28).
The outer surface (18a) of the wing section (12) manufactured by the casting process is machined according to the machining path to remove excess material to follow the generated target outer wing shape (40 _T). To do
Look at including it,
Finding the target outer wing shape (40 _{T) is}
After the casting process, the three-dimensional outer shape (40 _A ) of the blade section (12) was measured.
For the measured outer shape (40 _A ) of the cast blade section (12), a cooling passage position and shape measurement for the one or more internal cooling passages (28) is performed, and the cooling passage position and the cooling passage position and the shape are measured. shape measurement is performed by performing the actual wall thickness (T _a) measured at a plurality of points along said outer wall (18) of said cast airfoil section (12),
A point (42) representing the value _{of the nominal wall thickness (TN} ) around the measured position (28 m) of the one or more internal cooling passages (28) constitutes.
A best fit operation was performed to match the _{nominal outer wing shape (40 N} ) to the point (42) representing the value of the nominal wall thickness ( _TN).
The nominal outer wing shape after the best fit operation so as to coincide with a point (42a) representing a value of nominal wall thickness that still deviates from the nominal outer wing shape (40 _{N) matched by the best fit operation.} By adapting (40 _N ), the target outer wing shape (40 _T ) is generated.
Including that
Method. The target outer wing shape (40 _T ) so that the target outer wing shape (40 _T ) does not extend beyond the measured outer shape (40 _A ) of the cast wing section (12). further comprising the method of claim 1, to limit the. The method according to claim 1 or 2, wherein the measurement of the three-dimensional outer shape (40 _A ) of the wing section (12) is performed by tactile coordinate measuring machine probing, laser scanning or photogrammetry, or a combination thereof. The measurement of the actual wall thickness (T _A) is an ultrasound, X-rays, computed tomography is performed using an eddy current or a combination thereof, any one of claims 1 to 3 the method of. The actual wall thickness measurements (T _A), the carried out in the span direction (radial direction) and the chord various points along the direction of the cast airfoil section (12), The method of claim 4. A method of manufacturing a row of turbine blades or vanes,
Multiple turbine blades or vanes are manufactured by the casting process, each blade or vane having a blade section (12) with one or more internal cooling passages.
The outer surface (18a) of each blade section (12) is machined after the casting process by the method according to any one of claims 1-5.
A specific machining path used for the machining is generated in the wing section (12) of each blade or vane.
A method of manufacturing a row of turbine blades or vanes. A CAD module (56) for generating machining path data for adaptively machining a blade section (12) of a turbine blade or vane manufactured by a casting process, said blade section (12). It has an outer wall defining the inside of the wing with one or more internal cooling passages (28),
The CAD module (56) is configured to receive design data for the wing section (12), including _{nominal outer wing shape (40 N} ) and nominal wall thickness ( _{TN) data.}
The CAD module (56) is _{configured to generate the machined path data by determining the target outer blade shape (40 T} ), and the CAD module (56) is configured to generate the nominal wall thickness ( _TN). ) Is maintained at all points in the outer wall (18) around one or more internal cooling passages (28) in the wing section to be machined later. It is configured to generate the target outer wing shape (40 _{T ) by adapting N).}
The machined path data covers the outer surface (18a) of the wing section (12) manufactured by the casting process in order to remove excess material to follow the generated target outer wing shape (40 _T). It defines information for machining,
The CAD module (56) is _{configured to receive three-dimensional outer shape (40 A} ) measurement data for the cast wing section (12).
The CAD module (56) provides cooling passage position and shape measurements for one or more internal cooling passages (28) with respect _{to the measured outer shape (40 A) of the cast blade section (12).} is configured to perform, the cooling passage position and shape measurement, the measurement of the actual wall thickness (T _a) at a plurality of points along said outer wall (18) of said cast airfoil section (12) Made by doing,
The CAD module (56) constitutes a point (42) representing a value _{of nominal wall thickness (TN} ) around a measured position (28 m) of the one or more internal cooling passages (28). Adapted to
The CAD module (56) is adapted to perform a best fit operation to match the _{nominal outer wing shape (40 N} ) with respect to the point (42) representing the value of the _{nominal wall thickness (TN).} Have been made
The CAD module (56) is aligned with a point (42a) representing a nominal wall thickness ( _{TN ) value that still deviates from the nominal outer wing shape (40 N) matched by the best fit operation.} Is adapted to produce a target outer wing shape (40 _T ) by adapting the nominal outer wing shape (40 _{N) after the best fit operation.}
CAD module (56). Further, the CAD module (56) ensures that the target outer wing shape (40 _T ) does not extend beyond the _{measured outer shape (40 A} ) of the cast wing section (12). The CAD module (56) according to claim 7 , which is configured to limit the target outer wing shape (40 _T).

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